Spacecraft most commonly use photovoltaic solar cells to collect solar radiation and convert it into the electrical power necessary to operate the spacecraft. The solar cells are normally disposed on a solar array system. A solar array typically comprises one or more solar panels mechanically and electrically attached to each other and to the spacecraft. Each solar panel in an array typically comprises numerous individual photovoltaic solar cells, which are usually laid out in rows and connected together electrically at their adjacent edges. Most deployable solar arrays for spacecraft have used crystalline solar cells mounted to rigid honeycomb panels. Certain prior art describes mechanisms to effectively package, carefully deploy, and maintain the shape of rigid panel solar array designs. Flexible blanket solar arrays have also been used for crystalline solar cells, and are generally packaged in a long cylindrical roll or a foldable pleated stack that is deployed using separate or integrated mechanical structures.
Optimization of standard solar array systems, both structural and electrical subsystems, incorporating standard 140 um thickness multijunction solar cells with 100 um thick coverglass are approaching ˜80 W/kg BOL specific power and ˜13 kW/m3 stowed packaging performance plateaus. Current state-of-the-art optimized solar array systems utilize heavy carbon composite honeycomb panel structures to provide necessary deployed and stowed strength and stiffness to meet mission requirements. Other promising solar array arrays that incorporate flexible blanket technologies with high-efficiency multijunction solar cells promise even higher specific power beyond 100 W/kg and more compact stowed packaging performance beyond 30 kW/m3, allowing for more power to be packaged within a given launch volume envelope. These optimized structures and photovoltaic blanket subsystems are of less mass while stilling achieving typical stowed and deployed structural strength and stiffness performance. Furthermore, the solar cell panels on either a rigid panel or flexible blanket solar array construction account for nearly 80% of the cost of a space solar array system, and as such are a major target for cost reduction through the proposed embodiment.
To enable current and future spacecraft missions photovoltaic flexible blanket solar array technology must be affordable, modular, reliable, lightweight, accommodate standard and emerging photovoltaic solar cells, survivable and operable for over 15 years in multiple space environments (including thermal temperature extremes, proton and electron radiation, charged particle plasma environments, solar flares, atomic oxygen), operate at low 24 volt and high 600 volt voltages, low outgassing, stow within a compact volume for launch, provide dimensional stability, provide tensile load capability and long fatigue life, provide high strength and stiffness, provide insulating properties for the solar cell circuit, allow for good radiation and conduction properties to remove thermal heat.
The most notable flexible blanket solar arrays and their construction produced to date are: ATK's UltraFlex which is a radial rib structure with a tensioned single layer open weave mesh substrate blanket with an open backside circuit that is exposed to the environment; Lockheed Martin's Solar Array Flight Experiment (SAFE), International Space Station (ISS), and Milstar all of which are a central structure tensioned rectangular Kapton glass-reinforced laminated blanket; ESA/British Aerospace Hubble Space Telescope (HST) which is a dual-side structure tensioned rectangular Kapton glass-reinforced laminated blanket; AEG-Telefunken/Spar Aerospace L-SAT and Olympus which is a central structure with tensioned rectangular Kapton glass-reinforced laminated blanket; Northrop Grumman (TRW) EOS-AM/Terra and APSA which is a central structure with tensioned rectangular Kapton carbon-reinforced laminated blanket; and Boeing (Hughes) FRUSA which is a dual-sided structure tensioned rectangular Kapton glass-reinforced laminated blanket. With the exception of ATK's UltraFlex, all past solar array flexible blanket assemblies developed, built, and flown to date are constructed from a continuous area laminated frontside and backside polyimide/Kapton film with a heavy glass reinforcement core. Northrop Grumman's EOS-AM/Terra solar array implemented a lighter-weight carbon reinforcement core in place of the heavier glass material, but still employed a laminated construction. Northrop Grumman implemented carbon reinforcement within their blanket technology to further reduce mass and provide lower cell operating temperature through better thermal conductivity properties. ATK's UltraFlex flexible blanket assembly is very different and is constructed of a silicone coated glass fiber material where the solar cell circuit is directly bonded to the tensioned membrane glass fiber substrate, and the solar cell circuit is exposed on its backside with no dielectric for protection from the space environment. All these prior art flexible blankets are continuous area one-piece assemblies that are under significant tension on-orbit. The prior art one-piece blanket construction is not modular or reconfigurable and presents difficulty for rapid repair or re-configuration to a variety of planar aspect ratios. In the deployed configuration the prior art one-piece blanket assembly constructions are under significant tension on-orbit, and simultaneously the delicate solar cells and electrical interconnects are also in tension and not structurally isolated from the system. The fact that the delicate interconnected solar cells and circuits are in the load path presents a problem with respect to mechanical loading of the electrical subsystem which is not fundamentally designed to withstand significant loads. In prior art systems the interconnected solar cells, strings and circuits are exposed to additional mechanical loading which hinder reliability and can cause opening of electrical circuits. In addition to the aforementioned mechanical loading concerns within the electrical system, large thermal mismatches between the high thermal expansion polyimide based continuous blanket designs and the deployment structure are inherent drawbacks with these prior art flexible blanket solar arrays. In order to accommodate large thermal mismatches between the blanket and the structure complex tensioning mechanisms are required to prevent blanket and deployment structural mechanical failure. The proposed Integrated Modular Blanket Assembly (IMBA) embodiment eliminates these deficiencies associated with prior art flexible blanket technologies
U.S. Pat. No. 5,298,085 to Harvey used on ATK's UltraFlex solar array discloses a flexible photovoltaic panel composed of a blanket substrate comprised of glass fiber reinforced open weave cloth, whereby the cloth is coated in silicone where each solar cell is directly bond to the glass fibers through the silicone, and a large portion of the solar cell backside is exposed. The Harvey embodiment produces an exposed solar cell circuit which promotes unwanted power loss through parasitic current collection when operating in charged particle plasma environments. This deficiency degrades the total power output capability of the Harvey embodiment when operating within plasma environments. To form a barrier between circuits and prevent arcing from adjacent strings for high voltage operations, in either plasma or normal space environments, typically the industry applies silicone adhesive grouting between the solar cell edges of adjacent circuits. Silicone grouting techniques between solar cells is well documented and is a suitable survivability design provision that enables high-voltage operation and mitigation against arcing. The Harvey embodiment with its open backside surface does not allow for grouting, and as such is not an ideal design for high voltage operability and arc mitigation. Another deficiency with the Harvey embodiment is that because the individual solar cells are bonded directly to the fiber reinforced substrate, whereby the fiber reinforced substrate is the tension load bearing structure, the interconnected solar cell circuits become exposed to unwanted mechanical and thermal stresses during operation in the deployed state, which in turn greatly reduces the solar cell circuits fatigue life and more likely would promote an open circuit condition rendering power loss to the spacecraft. Yet another deficiency of the Harvey embodiment is the large amount of exposed silicone material that is coated on the panel assembly backside, which is problematic for the end-user and a major source of contamination from volatile outgassing. Finally, the Harvey embodiment is not modular and is not produced from standard power modules (SPM's), but rather each individual solar cell is integrated to the blanket assembly as an independent component and not a higher level assembly. This produces a rather labor intensive construction of the Harvey embodiment which does not promote affordability.
U.S. Pat. No. 4,755,231 to Kurland describes a single film Kapton/polyimide material construction flexible blanket solar panel assembly. The assembly has additional means formed integrally, such as coatings or paint, on the exposed Kapton material for providing a ground path for electric charge building up and for providing a suitable heat emissive surface on the panel backside. The Kurland embodiment provides many improvements over Harvey in that it encapsulates the backside of the solar cell circuit which eliminates current losses when operating in charged particle plasma environments. The Kurland embodiment also greatly minimizes exposed silicones significantly reducing outgassing volatiles. The weakness of the Kurland embodiment is that blanket tension loads are transferred through the delicate and low strength Kapton/polyimide film material. This material has been known to structurally degrade and tear over a mission life making the application of this material as a structural member for a tensioned membrane flexible blanket not ideal. As such, this material as the primary load bearing membrane has been known to tear and rip sometimes resulting in catastrophic failure. Additionally, the Kapton/polyimide material has very high coefficient of thermal expansion and as such produces rather large dimensional instability on-orbit. Finally, the Kurland embodiment is not modular and is not produced from standard power modules (SPM's), but rather each individual solar cell is integrated to the blanket assembly as an independent component and not a higher level assembly. This produces a rather labor intensive construction of the Kurland embodiment which does not promote affordability.
U.S. Pat. No. 4,968,372 to Maass is an improvement to Kurland and describes a reinforced glass fiber Kapton/polyimide laminated based substrate for flexible blanket solar panels. The primary improvement features of Maas versus Kurland are the laminated construction of the flexible panel assembly with glass fiber filaments to provide strength and tearing resiliency under on-orbit tension loads, and better matched coefficient of thermal expansion with respect to the solar cell circuit and the deployment structure. The Maass construction is a similar panel construction used in some past and present flexible blanket solar arrays such as the Olympus, NASA International Space Station, and Milstar flexible blanket solar arrays. Other variations similar to the Maass embodiment, such as the NASA EOS-AM/Terra flexible blanket solar array, have employed the integration of carbon fiber reinforcement laminated within the Kapton/polyimide stack. The carbon fiber reinforcement is an improvement to the glass fiber in terms of providing lower mass, higher strength, better matched coefficient of thermal expansion, and better thermal conduction performance. Finally, the Maas embodiment is not modular and is not produced from standard power modules (SPM's), but rather each individual solar cell is integrated to the blanket assembly as an independent component and not a higher level assembly. This produces a rather labor intensive construction of the Kurland embodiment which does not promote affordability.
Current and prior art space solar arrays and solar cell panels, either rigid or flexible blanket construction, account for nearly 80% of the cost of a space solar array system, and as such are a targeted area for cost reduction through the preferred embodiment. Current processes for solar cell panel construction involve very labor intensive operations and many solar cell panels produced for space missions are custom constructions of differing panel geometries, solar cell and string features, and technologies. Current and prior art solar panel designs are also not modular, making it difficult to mass produce with economical automation processes or even traditional labor intensive processes. Additionally, with current solar panel technologies if one part is imperfectly produced it affects a larger part of the entire solar array system which must be reworked at considerable cost and schedule to the program. Current solar array panel technology does not have modular construction and repair and rework are cumbersome and expensive. The preferred embodiment, which utilizes common mass-produced standard power modules (SPM's), attacks the current technologies deficiencies and greatly enhances affordability and reliability through modularity, rapid production, and rapid rework capability.